In the compressor portion of an aircraft gas turbine engine, atmospheric air is compressed to 10-25 times atmospheric pressure, and adiabatically heated to about 800°-1250° F. (425°-675° C.) in the process. This heated and compressed air is directed into a combustor, where it is mixed with fuel. The fuel is ignited, and the combustion process heats the gases to very high temperatures, in excess of 3000° F. (1650° C.). These hot gases pass through the turbine, where rotating turbine wheels extract energy to drive the fan and compressor of the engine, and the exhaust system, where the gases supply thrust to propel the aircraft. To improve the efficiency of operation of the aircraft engine, combustion temperatures have been raised. Of course, as the combustion temperature is raised, steps must be taken to prevent thermal degradation of the materials forming the flow path for these hot gases of combustion.
Every aircraft gas turbine engine has a so-called High Pressure Turbine (HPT) to drive its compressor. The HPT sits just behind the compressor in the engine layout and experiences the highest temperature and pressure levels (nominally 2400° F. and 300 psia respectively) developed in the engine. The HPT also operates at very high speeds (10,000 RPM for large turbofans, 50,000 for small helicopter engines). In order to meet life requirements at these levels of temperature and pressure, today's HPT components are always air-cooled and constructed from advanced alloys.
While a straight turbojet engine will usually have only one turbine (an HPT), most engines today are of the turbofan, either high bypass turbofan or low bypass turbofan, or turboprop type and require one (and sometimes two) additional turbine(s) to drive a fan or a gearbox. The additional turbines are called the Low Pressure Turbines (LPT) and immediately follows the HPT in the engine layout. Since substantial pressure drop occurs across the HPT, the LPT operates with a much less energetic fluid and will usually require several stages (usually up to six) to extract the available power.
One well-known solution that has been undertaken to protect the metals that form the flow path for the hot gases of combustion, including those of the HPT and LPT, have included application of protective layers having low thermal conductivity. These materials are applied as thermal barrier coating systems (TBCs), typically comprising a bond coat that improves adhesion of an overlying ceramic top coat, typically a stabilized zirconia, to the substrate. These systems are known to improve the thermal performance of the underlying metals that form the flow path in the hot section of the engine. However, as temperatures of combustion have increased, even these TBCs have been found to be insufficient.
Another solution that has been used in conjunction with TBCs is active cooling of metal parts. Initially, active cooling provided a flow of air from the compressor to the back side of metal parts comprising the flow gas path. As temperatures increased even further, serpentine passageways were formed in the metallic components and cooling air was circulated through the parts to provide additional cooling capability, the cooling air exiting through apertures positioned in the gas flow side of the component, providing an additional impingement film layer along the gas flow path. This method is referred to as film cooling. Even though the air from the compressor is adiabatically heated to perhaps as high as 1250° F. (675° C.), the compressor air is still significantly cooler than the combustion gases moving along the gas flow path of the engine. However, as the temperatures of the combustion process have continued to increase, even these tried and true methods of cooling are reaching their limitations. In particular, the turbine nozzles of high efficiency, advanced cycle turbine engines are prone to failure as a result of thermal degradation.
While some modifications of the traditional flow path surfaces have been applied in the past, such as the application of materials over the TBC, these modifications have been directed to reducing the emissions of pollutants such as unburned hydrocarbons (UHC) and carbon monoxide (CO). One such modification is set forth in U.S. Pat. No. 5,355,668 to Weil, et al., assigned to the assignee of the present invention, which teaches the application of a catalyst such as platinum, nickel oxide, chromium oxide or cobalt oxide directly over the flow path surface of the thermal barrier coating of a component such as a turbine nozzle. The catalyst layer is applied to selected portions of flow path surfaces to catalyze combustion of fuel. The catalytic material is chosen to reduce air pollutants such as unburned hydrocarbons (UHC) and carbon monoxide (CO) resulting from the combustion process. The catalytic layer is applied to a thickness of 0.001 to 0.010 inches and is somewhat rough and porous, having a surface roughness of about 100 to 250 micro inches, in order to enhance the surface area available to maximize contact with the hot gases in order to promote the catalytic reaction. The rough surface assists in creating some turbulence that promotes contact with the catalytic surface.
The prior art solutions are either directed to problems that are unrelated to the problem of high temperature experienced by turbine nozzles, such as the Weil patent, or provide different solutions to the problem of high temperatures resulting from the combustion process. The present invention provides a different approach to the problem of high temperatures experienced by turbine nozzles.